JP2010203439A - Effusion cooled one-piece can combustor - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a means for cooling a component of a gas turbine and effusion cooling of an one-piece can combustor. <P>SOLUTION: A combustor for an industrial turbine includes a single transition piece transitioning directly from a combustor head end to a turbine inlet. The transition piece includes an inner surface and an outer surface. The inner surface bounds an interior space for combusted gas flow from the combustor head end to the turbine inlet. The outer surface at least partially defines an area for compressor discharge air flow. The transition piece includes a plurality of apertures configured to allow the compressor discharge air to flow into the interior space. Each of the plurality of apertures extends from an inlet portion on the outer surface to an outlet portion on the inner surface. <P>COPYRIGHT: (C)2010,JPO&INPIT

Description

  The present invention relates generally to means for cooling gas turbine components, and more specifically to jet cooling of a single component can combustor.

  A gas turbine can operate very efficiently if the turbine inlet temperature can be raised to a maximum value. However, the combustion chambers from which combustion gases are generated and flow into the turbine inlet reach operating temperatures well above 1500 ° F., even with the most advanced alloys over long periods of use. Cannot withstand. Thus, turbine performance and lifetime are highly dependent on the degree of cooling that can be provided to turbine components that are exposed to extreme heating conditions.

US Pat. No. 7,082,766

  The general concept of cooling turbine components using compressor discharge air is known in the art. However, turbine design developments and variations do not necessarily involve a specific structure in which a cooling mechanism for turbine components is implemented. Accordingly, there is a need to implement a cooling mechanism in a newly developed turbine design.

  The following presents a simplified summary of the invention in order to provide a basic understanding of some exemplary aspects of the invention. This summary is not an extensive overview of the invention. Furthermore, this summary is not intended to identify key elements of the invention or to accurately describe the scope of the invention. Its sole purpose is to present some concepts of the invention in a simplified form as a prelude to the more detailed description that is presented later.

  To achieve the above and other aspects in accordance with the present invention, a can combustor for an industrial turbine is provided that includes a single transition piece that transitions directly from the combustor head end to the turbine inlet. The transition piece defines an external space for the compressor discharge air flow and an internal space for the combustion gas flow. The transition piece includes an outer surface that defines an outer space boundary and an inner surface that defines an inner space boundary. The transition piece includes a plurality of apertures configured to allow the compressor discharge air flow to flow into the interior space. Each of the plurality of apertures extends from an inlet portion on the outer surface to an outlet portion on the inner surface.

  According to another aspect of the invention, an industrial turbine engine includes a combustion section, an air discharge section downstream of the combustion section, a transition region between the combustion section and the air discharge section, and a combustor transition component. Including. The combustor transition piece defines a combustion section and a transition region. The transition piece is adapted to carry the combustion gas stream to the first stage of the turbine corresponding to the air discharge section and defines an external space for the compressor discharge air stream and an internal space for the combustion gas stream. The transition piece includes an outer surface that defines an outer space boundary and an inner surface that defines an inner space boundary, and includes a plurality of apertures configured to allow compressor discharge airflow to flow into the inner space. Each of the plurality of apertures extends from an inlet portion on the outer surface to an outlet portion on the inner surface.

1 is a schematic cross-sectional view of an exemplary embodiment of a single configuration can combustor in which the present invention can be practiced. The expansion perspective view of the transition component provided with the jet hole. Sectional drawing which crosses the jet hole of a transition component.

  The foregoing and other aspects of the present invention will become apparent to those skilled in the art to which the present invention relates upon reading the following description with reference to the accompanying drawings.

  Illustrative embodiments incorporating one or more aspects of the invention are described and illustrated in the drawings. These illustrated embodiments are not intended to limit the invention. For example, one or more aspects of the present invention can be utilized in other embodiments and can also be utilized in other types of devices.

  FIG. 1 illustrates an embodiment of a single configuration combustor 10 in which the present invention may be implemented. The illustrated exemplary embodiment is a can-type annular backflow combustor 10, but the invention is applicable to other types of combustors. The combustor 10 burns air and fuel in a limited space, and discharges the resulting combustion gas through a fixed train of vanes, thereby providing the gas necessary to drive the rotational motion of the turbine. appear. During operation, the discharge air from the compressor reverses direction as it passes the outside of the combustor 10 and flows again into the combustor 10 on the way to the turbine. Pressurized air and fuel are combusted in the combustion chamber. Combustion gas enters the turbine section at high speed via the transition piece 120. As discharge air flows over the outer surface of the transition piece 120, it provides convective cooling to the combustor components.

  In FIG. 1, the transition piece 120 transitions directly from the circular combustor head end 100 to the turbine annulus sector 102 (corresponding to the first stage of the turbine, indicated at 16) in a single configuration. The single component transition piece 120 can be formed from two halves or multiple components that are welded or joined together to facilitate assembly or manufacture. The sleeve 129 also transitions directly from the circular combustor head end 100 to the rear frame 128 of the transition piece 120 in a single configuration. The single-piece sleeve 129 is formed from two halves and can be welded or joined together to facilitate assembly. The joint between the sleeve 129 and the rear frame 128 forms a substantially closed end with respect to the cooling annulus 124. Also, “single” means a plurality of parts and / or unitary structures and / or a single configuration, and the like, joined together based on any suitable means for joining the elements. .

  In FIG. 1, there is an annular flow of discharged air that has been convectively treated over the outer surface of the transition piece 120. In an exemplary embodiment, the discharge air flows through a sleeve 129 that forms an annular gap that allows the flow rate to be sufficiently high to produce a high heat transfer coefficient. A sleeve 129 surrounds the transition piece 120 and forms a flow annulus 124 therebetween. The cross-flow cooling air moving in the annulus 124 continues to flow upstream as indicated by arrows. In an alternative embodiment, the sleeve 129 cannot extend completely from the combustor head end 100 to the rear frame 128. The circular area of the transition piece 120 will be discussed in more detail in FIGS.

  In conventional combustors, the combustor liner and flow sleeve are generally found upstream of the transition piece and sleeve, respectively. However, in the single configuration can combustor of FIG. 1, the combustor liner and flow sleeve have been eliminated to provide a shorter length combustor. The major components of a single component can combustor include a circular cap 134, an end cover 136 that supports a plurality of fuel nozzles 138, a transition piece 120, and a sleeve 129, which are known in the art. . For example, a more detailed description of a single configuration can combustor can be found in US Pat. No. 7,082,766 issued to Widener et al.

  FIG. 2 illustrates an embodiment of a single configuration transition piece 120 formed with a plurality of apertures or jets 200 in a separated state. FIG. 2 shows one exemplary arrangement of apertures 200 near the combustor head end 100 for ease of explanation only, which is not to be construed as a limitation of the present invention. Note that this is not possible. Thus, the formation of the aperture 200 can be in another selected region of the transition piece 120, or can extend to that region, or can span the entire outer surface. The selected area in which the aperture 200 is formed tends to be relatively hotter than other areas during turbine operation, and thus can be a spot on the transition piece 120 that can benefit from further cooling. It can be. Alternatively, the apertures 200 can be formed in a circumferentially distributed manner or can extend from an upstream portion to a downstream portion of the transition piece 120. Further, FIG. 2 shows only one of a plurality of possible arrangements in which a plurality of apertures 200 can be patterned. For example, the apertures 200 can be arranged orthogonally around each other. In another embodiment, each aperture 200 in a row can be slightly offset relative to the aperture in the adjacent row. Such diversity of sequences is within the scope of the present invention.

  FIG. 3 shows a cross-sectional view through an aperture 200 formed through a wall 300 that is part of the transition piece 120. Again, a limited number of apertures 200 are shown on the transition piece 120 for ease of explanation. FIG. 3 shows the outer surface 300 a and the inner surface 300 b of the wall 300. The area on the wall is the external space 302 of the transition piece 120, and the area below the wall is the internal space 304 of the transition piece 120. As described above, depending on the embodiment or portion of the transition piece 120, the sleeve 129 may or may not be adjacent to the transition piece 120, so that the flow annulus 124 is in this region. It may or may not be formed. When sleeve 129 is present, sleeve 129 is part of outer space 302 and flow annulus 124 will be formed between sleeve 129 and transition piece 120.

  The right side of FIG. 3 corresponds to the upstream region of the turbine, and the left side of FIG. 3 corresponds to the downstream region of the turbine. Thus, a flow H composed of hot gas is generated from the combustion chamber and is directed downstream of the interior space 304 of the transition piece 120. The flow C composed of compressed discharge air having a temperature lower than that of the combustion hot gas is generated from the compressor, but approaches the transition part 120 from the downstream region of the turbine and moves to the upstream side of the outer space 302 of the transition part 120. This is the same as a normal can-type annular backflow combustor.

  The aperture 200 extends from the outer surface 300a of the wall 300 to the inner surface 300b. The present invention includes an aperture 200 formed perpendicular to the wall 300 and an aperture 200 formed in the wall 300 at an angle θ. In FIG. 3, the aperture 200 is shown at an angle θ such that the outlet portion 200b of the aperture 200 is downstream or rearward with respect to the inlet portion 200a of the aperture 200. In one embodiment, the angle θ is formed by a longitudinal axis 200c of the aperture 200 and a direction 202 that is tangential to the wall 300 and directed downstream. The angle θ can be an acute angle of 30 degrees and can be in the range of 20 to 35 degrees. However, other smaller or larger angles are also contemplated. In FIG. 3, the downstream tangent points to the left. The second aperture 200 is substantially cylindrical, but if the aperture 200 is not perpendicular to the wall 300, the inlet portion 200a and the outlet portion 200b will have an elliptical shape. However, the apertures 200 and 400 are not circular, and may have a cross section such as a polygon.

  Another variation of the aperture 200 is that the angular position of the inlet portion 200 a can be different from the angular position of the outlet portion 200 b on the circumference of the transition piece 120. Further, the outlet portion 200b of the aperture 200 can be upstream or forward with respect to the inlet portion 200a of the aperture 200, thereby creating an obtuse angle between the longitudinal axis of the aperture 200 and the direction 202.

  In FIG. 3, the aperture 200 has a generally cylindrical geometry with a constant diameter from the inlet portion to the outlet portion. In one embodiment, the diameter can be 0.03 inches or in the range of 0.02 inches to 0.04 inches. Of course, other diameters of the aperture 200 are also contemplated. For example, the aperture 200 can gradually increase or decrease in diameter through the wall 300.

  The aperture 200 can be formed through the wall 300 of the transition piece 120 by laser drilling or other machining methods selected based on factors such as cost and accuracy.

  In FIG. 3, stream C provides convective cooling of the transition piece 120 by removing heat while passing the outer surface 300a. The flow E generated by the aperture or jet hole 200 provides an air jet in the entire or selected area of the transition piece 120, which passes through the aperture 200 where it contacts the inner surface. Cool down. Jet cooling is a form of leaching cooling. Apertures other than those perpendicular to the wall 300 have a larger inner surface compared to the apertures perpendicular to the wall due to the increased length so that heat transfer is extended and more of the transition piece 120 Great cooling can be achieved. Further, after cooling air exits the exit portion 200b of the aperture 200, a layer or film of cooling air is formed adjacent to the inner surface 300b of the wall 300 of the transition piece 120. Formation of such a layer of cooling air on the inner surface 300b further cools the transition piece 120. Since the degree of change in direction required by the cooling air is reduced, the formation of such a layer is facilitated by the angled aperture compared to the right angle aperture. However, the present invention includes two variations of right angle and angled apertures. Cooling by the film formed on the inner surface can be improved as the pore size and angle are reduced. However, the smaller the pores, the stronger the tendency to block impurities. In comparison, larger holes can cause excessive penetration of the hot gas stream by the cooling air jet, reducing turbine efficiency. Accordingly, such advantages and disadvantages need to be considered overall when determining the geometry of the jet holes.

  The invention has been described with reference to the exemplary embodiments described above. Modifications and alternatives will occur to others upon reading and understanding this specification. Exemplary embodiments incorporating one or more aspects of the invention are intended to include all such modifications and alternatives as long as they are within the scope of the appended claims.

DESCRIPTION OF SYMBOLS 10 Single component combustor 10 Can-type annular backflow combustor 10 Combustor
10 Combustor 100 Circular Combustor Head End (2)
100 Combustor head end (2)
102 Turbine annulus sector 120 Transition parts (29)
120 single component transition part 120 single component transition part 124 cooling annulus 124 flow annulus (3)
124 Annulus 128 Rear frame (3)
129 sleeve (10)
129 Single configuration sleeve 134 Circular cap 136 End cover 138 Multiple fuel nozzles 200 Jet holes (2)
200 Aperture (23)
200 Aperture Selection Area 200 Multiple Apertures 200 Each Aperture 200 A Limited Number of Apertures 200 200 Including Apertures Second Aperture 200a Inlet Portion (2)
200a Entrance part (2)
200b Exit part (2)
200b Exit part (3)
200c Longitudinal axis 202 direction (2)
300 walls (11)
300a Outer surface (3)
300b inner surface (4)
302 External space (3)
304 Internal space (2)

Claims (15)

An industrial turbine combustor with a single transition piece (120) that transitions directly from the combustor head end (100) to the turbine inlet (16),
The transition piece (120) includes an inner surface (300b) and an outer surface (300a), the inner surface (300b) burning from the combustor head end (100) to the turbine inlet (16). Demarcating an interior space (304) for gas flow, the outer surface (300a) at least partially defining a region (302) for compressor discharge air flow (C), and the transition piece comprises: A plurality of apertures (200) configured to allow the compressor discharge air stream to flow into the interior space (304), each of the plurality of apertures (200) being on the outer surface (300a); A combustor extending from an inlet portion (200a) to an outlet portion (200b) on the inner surface (300b). One of the inlet part (200a) and the outlet part (200b) is disposed further downstream than the other of the inlet part (200a) and the outlet part (200b).
The combustor according to claim 1. The combustor is in a direction in which the combustion gas flow (H) and the compressor discharge air flow (C) face each other, and a longitudinal axis (200c) passing through the aperture (200) is a direction of the combustion gas flow (H). A can-type annular backflow type configured to form an acute angle (θ) with respect to the direction of the compressor discharge air flow (C).
The combustor according to claim 2. The longitudinal axis (200c) passing through the aperture (200) is oriented to form an acute angle (θ) with respect to the downstream tangent of the outer surface (300a);
The combustor according to claim 1. The acute angle (θ) is in the range of 20 ° to 35 °,
The combustor according to claim 4. The plurality of apertures (200) have a constant diameter ranging from 0.02 inches to 0.04 inches from the inlet portion (200a) to the outlet portion (200b);
The combustor according to claim 1.   The combustor of any preceding claim, wherein the aperture (200) is substantially perpendicular to the outer surface (300a). The transition piece (120) is unjoined;
The combustor according to claim 1. A combustion section;
An air discharge section downstream of the combustion section;
A transition region between the combustion section and the air discharge section;
A combustor transition piece (120) defining the combustion section and the transition region;
In industrial turbine engines including
The transition piece (120) is adapted to carry a combustion gas stream (H) to a first stage of a turbine corresponding to the air discharge section;
The transition piece (120) includes an inner surface (300b) and an outer surface (300a);
The inner surface (300b) delimits the internal space (304) of the combustion gas stream (H) from the combustor head end to the turbine inlet;
The outer surface (300a) at least partially defines a region (302) for the compressor discharge airflow (C);
The transition piece (120) includes a plurality of apertures (200) configured to allow the compressor discharge airflow (C) to flow into the interior space (304);
Each of the plurality of apertures (200) extends from an inlet portion (200a) on the outer surface (300a) to an outlet portion (200b) on the inner surface (300b).
Industrial turbine engine. One of the inlet part (200a) and the outlet part (200b) is disposed further downstream than the other of the inlet part (200a) and the outlet part (200b).
The industrial turbine engine according to claim 9. The combustor transition component (120) is in a direction in which the combustion gas flow (H) and the compressor discharge air flow (C) face each other, and a longitudinal axis (200c) passing through the aperture (200) is the combustion gas flow. It is a can-type annular backflow type configured to form an acute angle (θ) with respect to the direction of (H) and to form an obtuse angle with respect to the direction of the compressor discharge air flow (C).
The industrial turbine engine according to claim 10. The longitudinal axis (200c) passing through the aperture (200) is oriented to form an acute angle (θ) with respect to the downstream tangent of the outer surface (300a);
The industrial turbine engine according to claim 9. The acute angle (θ) is in the range of 20 ° to 35 °,
The industrial turbine engine according to claim 12. The plurality of apertures (200) have a constant diameter ranging from 0.02 inches to 0.04 inches from the inlet portion (200a) to the outlet portion (200b);
The industrial turbine engine according to claim 9. The aperture (200) is substantially perpendicular to the outer surface (300a);
The industrial turbine engine according to claim 9.

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